This invention relates to turbomachinery blades and particularly to a blade having a unique suction surface contour that mitigates shock induced aerodynamic losses.
Gas turbine engines and similar turbomachines employ a turbine to extract energy from a stream of working medium fluid. A typical axial flow turbine includes one or more arrays of blades that project radially from a rotatable hub. The blades circumferentially bound a series of interblade fluid flow passages. Under some operating conditions, the working medium may accelerate to a supersonic speed as it flows through the interblade passages. The fluid acceleration produces expansion waves; subsequent deceleration produces compression waves and an accompanying primary shock that originate near the trailing edge of each blade and extend across the passage to the suction surface of the neighboring blade. A secondary or xe2x80x9creflectedxe2x80x9d shock, related to the primary shock, may also develop. The secondary shock extends into the working medium fluid stream downstream of the blade array.
The shocks degrade turbine efficiency by causing an unrecoverable loss of the fluid stream""s stagnation pressure. The shocks also interact with the fluid boundary layer attached to the suction surfaces of the blades, causing the boundary layer to enlarge and thereby introducing additional aerodynamic inefficiencies. The shocks also introduce static pressure pulses into the fluid stream. These pressure pulses impinge upon turbine components downstream of the blade array and subject those components to increased risk of high frequency fatigue failure. Clearly, it is desirable to eliminate or mitigate these adverse effects of the shocks to ensure peak turbine efficiency and to enhance the durability of the turbine components.
It is, therefore, a principal object of the invention to provide a turbomachinery blade that influences the pattern of expansion waves and shocks in a way that weakens or eliminates the shocks.
According to one aspect of the invention, the airfoil of a turbomachinery blade has a uniquely contoured suction surface with chordwisely separated, positively curved forward and aft segments and a negatively curved medial segment residing chordwisely intermediate the positively curved segments. The medial segment may extend across substantially the entire span of the blade or may be spanwisely localized. When used in a turbomachinery blade array, the medial segment limits expansion of the fluid stream as it accelerates through the passages. Consequently, the degree to which a shock must subsequently recompress and decelerate the fluid stream to satisfy the aerodynamic boundary conditions imposed on the fluid stream is similarly limited. As a result, the primary and secondary shocks are weaker and therefore less detrimental to turbine efficiency. Under some conditions, the secondary shock may not even materialize.
The principal advantage of the invention is the improved efficiency arising from reduced aerodynamic losses. A related advantage is the reduced risk of exposing the turbine components to premature high frequency fatigue failure.
The foregoing objects and advantages and the operation of the invention will become more apparent in light of the following description of the best mode for carrying out the invention and the accompanying drawings.